Generally described, a turbine stage of a gas turbine engine includes a number of stationary turbine nozzles. Each turbine nozzle may have a vane extending radially between outer and inner sidewalls. The nozzle vanes may have an airfoil configuration for guiding the combustion gases between corresponding turbine rotor blades disposed upstream and downstream thereof. The turbine rotor blades may be mounted to the perimeter of a rotor disk for rotation therewith. Because the turbine nozzle vanes are heated during operation by the hot combustion gases that flow therethrough, cooling air bled from the compressor may be channeled inside the vanes for cooling purposes. Limiting the amount of parasitic cooling air required and limiting the leakage of such cooling air lost in the nozzle vanes and elsewhere should promote overall gas turbine engine efficiency and performance.
Compartmentalized cooling has been used in the past with aviation turbine engines and the like. Such aviation engines generally include a circular (360°) component to direct the cooling flow into the nozzles. This configuration may be possible with aviation engines given that aviation engines generally are full hoop case structures that are axially stacked during assembly. Due to the overall size of industrial gas turbine engines, however, such industrial gas turbines generally are installed in at least two half (180°) segments, if not many more. This segmented configuration generally precludes the use of a 360° component to route the cooling flow into the nozzle arrangement.
There is thus a desire for an improved industrial gas turbine design. Such an improved industrial gas turbine design may use a number of segmented cooling baffles to provide high pressure cooling air with low leakage so as to promote efficient cooling with low leakage.